This invention relates to a recirculation structure for turbo-compressors, a turbo-compressor, an aircraft engine, and a stationary gas turbine.
Recirculation structures for turbo-compressors have been known in the art for quite a long time, and in the trade are generally referred to as “casing treatments”. Their primary function is to increase the aerodynamically stable operating range of the compressor, wherein the so-called surge margin is shifted to higher compressor pressures, i.e. to a higher compressor load. The failures that are responsible for localized stalling and ultimately for the surging of the compressor occur on the casing side at the ends of the rotating blades of one or more compressor stages, and on the hub side at the ends of the vanes that lie radially inside, because in these areas the aerodynamic load is the highest. By recirculating the “air particles” that circulate between the blade tips at blade speed, and whose energy level is reduced, into the main stream with an increase in energy, the flow in the area of the blade ends is again stabilized. Because flow disruptions as a rule do not occur evenly over all the stages, in terms of fluid mechanics, a circumferential balancing, in addition to the essentially axial recirculation, should also be possible. The primary disadvantage of known “casing treatments” is that, although they do increase the surge margin, they also reduce the efficiency level of the compressor.
German publication DE 33 22 295 C3 protects an axial fan with a “casing treatment”. Recognizable therein is an annular chamber (8) in which guide vanes (9) are fixed. In the downstream area over the ends of the rotating blades is an area that is open circumferentially into which the guide vanes do not extend. Characteristic of this type of “casing treatment” is a closed ring (7) that is aligned generally with the shape of the main flow channel, with the ring separating the rear intake area from the forward outlet area of the recirculation structure, and forming a smooth, closed surface area.
A quite similar “casing treatment” is known from German publication DE 35 39 604 C1, wherein an area that is open circumferentially is present in the forward and rear areas of the annular chamber (7). The radially inside ring 6 is also seen here.
A more recent “casing treatment” is known from U.S. Pat. No. 5,282,718. Here, the annular chamber (18, 28) and the guide vanes (24) are improved in terms of fluid mechanics. Here again, the intake and outlet of the recirculation flow are separated by a solid ring that is smooth and closed on the side of the blades. Rings of this type in the blade area must ordinarily be provided with a contact or intake coating in case they should come in contact with the blade tips.
Further “casing treatments” having axial or axially angled grooves are disclosed e.g. in U.S. Pat. No. 5,137,419. These are not taken into consideration here, however, because the grooves are not connected to one another in these versions; hence no circumferential flow comparison is possible.
U.S. Pat. No. 4,511,308 relates to ventilators (fans, blowers) having various “casing treatment” designs. The simplest design according to FIG. 6 possesses only one annular chamber without guide vanes. In the embodiments according to FIGS. 1 and 3, guide vanes are mounted in the annular chamber, and the upstream casing wall (22) is extended beyond the radial inside edges of the guide vanes (21) like a cylindrical or conical socket piece, so that on the upstream, forward end of the annular chamber no outlet of the recirculation flow into the main flow is possible. FIG. 5 shows guide vanes (21) that are mounted on the forward end wall and on the outer circumference of the annular chamber, and are further designed to be freestanding. Here there is no socket- or ring-like element that connects or covers the guide vanes circumferentially. The free, radially inside edges of the guide vanes (21) rise from the front to the back from the diameter of the intake casing (15) up to the greatest diameter of the annular chamber (16). In this manner, although in the downstream area the guide vanes axially overlap the upstream area of the blade ends (14), due to the great radial distance between the guide vanes (21) and the blade ends (14), no effective and defined guidance of the recirculation air is possible. Of further disadvantage is the large volume of the annular chamber (16) in relation to the blade dimensions. This type of embodiment is neither aerodynamically nor constructively suitable for a turbo-compressor.
In view of the disadvantages of the state of the art solutions, the object of the invention is to provide a recirculation structure for turbo-compressors that will enable a substantial enhancement of the surge margin, and thus a clear expansion of the stable operating range, without a significant reduction of the efficiency of the compressor.
The essence of the invention is to have the tips of the guide vanes that face the annular chamber lie on or near the contour of the main flow channel, and axially overlap the free ends of the blades or axially border the area of the free ends of the blades. In this manner, annular elements with contact coatings, etc. can be eliminated. The above-cited patent specifications show that up to now the professional world has consistently tried to design recirculation structures that will be smooth, without gaps, and closed over the largest possible axial range, up to the main flow channel, i.e. up to the so-called annular chamber, in order to effect an extension of the contours of the main flow channel that will be as favorable in terms of flow and loss as possible. The invention, in contrast, leads to gaps, fissured surfaces, etc., and thus would appear to be disadvantageous and inexpedient. Test have shown, however, that the recirculation structure of the invention is superior to known solutions in terms of both enhancing the surge margin and improving the level of efficiency. This can be explained in aerodynamic terms in that the free, informal design of the recirculation flow in the open annular chamber with free-standing guide vanes and circumferential flow links is more important than the greatest possible gap-free extension of the contour of the main flow channel. The absence of a closed ring has the further advantages that no contact or intake coating of the guide vanes is necessary, and radial space and weight are saved, resulting in structural mechanical advantages. However, a defined control of the recirculation flow—without annular elements—is achieved only if the free edges of the guide vanes run relatively close to the edges of the blades, partially overlapping them axially, or at least lying adjacent to their space. Only in this way can a compact “casing treatment” that is suitable for use in a compressor ultimately be achieved.
The ratio of the axial length of the annular chamber to the axial length of the blade ends preferably is 0.2 to 1.5. In the case of wide blades having a large axial span at the blade end, the ratio will be closer to 0.2; with narrow blades having a small axial span at the blade end, the ratio will be closer to 1.5.
In the preferred embodiment, the ratio of the radial height to the axial length of the annular chamber is 0.1 to 1.0. With aircraft engines having very strict standards in terms of space requirements, contours, etc., attempts will be made to manage with a smaller ratio, i.e. a smaller radial height. For stationary applications having adequate space available it is possible to go closer to the upper limit. With axially short annular chambers one would also tend more to approach the upper limit.
It is further preferred that the tips of the guide vanes that face toward the annular chamber are radially recessed, at least in the area of the free blade ends, far enough that during normal operation the ends of the blades will not come into contact with the guide vanes. This is due to the fact that the tips of the blades may be damaged by brushing against something, especially against hard, inflexible guide vane tips. The recessing of the guide vane tips is not in contradiction to the requirement that the tips should lie on or near the contour of the main flow channel, because the small radial gap dimensions required to prevent contact are practically without consequence in terms of fluid mechanics, i.e. they do not negatively affect recirculation.
Preferred embodiments of the recirculation structure, as well as a turbo-compressor, an aircraft gas turbine, and a stationary gas turbine all form the subject matter of the invention.